In the field of aircraft construction the classic aluminum materials are increasingly displaced by the use of fiber-reinforced composite materials that comprise, for example, carbon-fiber-reinforced duroplastic or thermoplastic synthetic materials. Often already complex structural components such as flaps or entire vertical stabilizers are manufactured throughout using such fiber-reinforced composite materials, in particular using CFRP materials.
Because of their spatial dimensions or their complex geometric shape such aerodynamically formed aircraft components are, as a rule, made in the so-called differential construction, in which the aircraft components need to be assembled, in a final installation step, from a multitude of prefabricated individual components of simpler geometric shapes.
In this context a landing flap for an aircraft is mentioned as an example of the aforesaid, in which landing flap on several spars that are spaced apart from each other so as to be parallel and that extend in longitudinal direction, a multitude of ribs, which extend across the aforesaid, are fastened to support the outer skin. Lastly, the surface geometry of the aircraft component is defined by the outer contour of the ribs and the shape of the outer skin, and thus the aerodynamic behavior of the landing flap is determined. Moreover, all the components need to be installed in a strain-free manner in order to avoid introducing additional loads into the structure.
This differential construction is associated, among other things, with one disadvantage in that the individual parts need to be joined in an additional installation step to form the finished component. Furthermore, the connection process generally requires overlaps or flanges between the individual components, which overlaps or flanges are associated with corresponding additional weight of the aircraft component.
Further disadvantages arise as a result of the riveted joint of the individual components, which riveted joint is used as a rule. Because fiber-reinforced composite components have considerably lower strengths of the hole walls when compared to those of metallic materials, each rivet hole represents a disadvantage in terms of statics, which disadvantage needs to be compensated by greater materials thicknesses in the region of the hole. To make the use of such riveted joints possible at all on fiber-reinforced composite components, it is necessary, for example on shell structures, to also provide greater material thicknesses and enlarged flange regions so that in the case of failure of the riveted joint it is indeed possible to effect repair work by creating a further connection. All these limitations result in the aircraft component not being designed with regard to the maximum mechanical load to be expected, but instead with regard to manufacturing constraints, an approach which unnecessarily results in greater weight.
In principle the structural components of an aircraft component can also be joined by bonding, whereby at least the problem of the reduced strength of the hole walls is eliminated. However, the so-called structural bonding of highly loaded components of an aircraft is still associated with considerable problems in terms of the required surface pre-treatment, fatigue resistance, and resistance to impact loads, which problems, for safety reasons, at present do not yet make possible its use in the field of civil aviation.
The so-called integral construction represents a feasible alternative to the differential construction; in integral construction fiber-reinforced composite components of complex geometric shape are made in a single piece so that the above-mentioned disadvantages resulting from connecting a multitude of individual parts to form a complex overall structure do not apply.
One problem in the manufacture of such aircraft components, which can, for example be complete flaps, airbrakes, ailerons, flap tracks, slats, engine mounts, winglets, wings or airfoils, tail units, control surfaces and the like, is associated with the undercut structures, which in many cases are necessary to create the necessary stiffening reinforcement within the closed outer skin. These undercut structures can be skin reinforcements and spar reinforcements that extend into the interior of the essentially hollow aircraft component.
From DE 10 2008 013 759 A1 a technical solution to this undercut problem in the manufacturing of fiber-reinforced aircraft components in integral construction is known. To make industrial production of such aircraft components possible requires mold cores that after manufacture of the aircraft component can be removed from said component without any resistance. It has been proposed to prefabricate mold cores from a soluble material in a core shape. In this arrangement these mold cores reproduce the inner surface geometry of the aircraft component. Subsequently, preforms comprising reinforcement fibers are placed onto these mold cores in order to form stiffening elements and to arrange the mold cores to form an overall structure. Subsequently a web-shaped semifinished product is placed onto the mold cores in order to create the outer skin. The overall structure is then placed in a closed mold tool and is infiltrated by a curable plastic matrix. This infiltration process, which is known per se, is also referred to as an RTM process. After the overall structure has cured to become the finished fiber-reinforced composite component with the application of pressure and temperature, the mold cores are removed by releasing them from the fiber-reinforced composite component.
While with this approach it is possible to control undercut contours in aircraft components of the type of interest in the present context, it appears, however, that manufacture of the soluble mold cores is quite elaborate and said mold cores can only be used once.
Apart from the above, from the generally-known state of the art multi-part mold cores are known that comprise a solid material, for example an aluminum alloy, and that divide the cross section of the mold core into at least nine core parts per bay. In the case of a nine-piece mold core, in a cross section of a bay, which cross section is rectangular as a rule, there is thus a center core part which after curing of the aircraft component needs to be pulled from the bay first, before the remaining core parts, which in each case are in contact with the outer skin or with the spars, can be removed, in that they are first pushed in the direction of the removed center core part in order to overcome undercuts, and are subsequently pulled longitudinally from the bay.
As a result of the large number of core parts the expenditure for cleaning the mold core comprising said core parts increases. Furthermore, in the case of trapezoidally tapering flaps the cross section of the individual core parts becomes quite thin, which makes handling said core parts more difficult when manufacturing the aircraft component. Such thin core parts can easily bend.
Furthermore, multi-part mold cores comprising only three core parts are generally known; however, their application is limited to bays of aircraft components in which only the outer skin or only the spar comprises inwards-directed undercuts. Thus their field of application is correspondingly limited.